This invention relates to gas turbine engines and, more particularly, to variable area exhaust nozzles for use therein.
The invention herein described was made in the course of or under a contract, or a subcontract thereunder, with the United States Department of the Air Force.
In modern, high-speed, multimission aircraft a requirement frequently exists for the gas turbine engines to have an augmented thrust (afterburning) capability. This necessitates large nozzle exit area variations as the engine switches between the augmented and non-augmented modes. In the subsonic, non-augmented mode the nozzle exit area is considerably smaller than the maximum engine diameter. Whether the engines are mounted within pods or within the aircraft structure (wing or fuselage), rapid external cross-sectional area closure rates are experienced. In other words, the external flowpath of the engine converges rapidly to the nozzle exit. These high area closure rates quite often result in high pressure drag.
Pressure drag penalties are incurred because air flowing along the aircraft surfaces, for example, may wholly or partially separate from the afterbody (or convergent) surfaces due to rapid contour changes associated with the aforementioned high closure rate. Separation of flow will usually create reduced pressures (below ambient) on the afterbody surfaces with a resultant increase in aircraft drag. A region particularly vulnerable is that between round axisymmetric nozzles of laterally adjacent engines operating at non-augmented power levels. This region is particularly difficult to fill with high energy ambient air because of the diverging nature of the space between nozzles and the complex flow field involved. Efforts to fill this space with mechanical structure to improve the area distribution have been only marginally successful due to the variable area requirements for the round axisymmetric exhaust nozzles.
One solution to this problem is to contour the exhaust nozzle such that it conformably mates with the adjacent structure, the preferred profile being asymmetric rather than axisymmetric, and a convenient shape being substantially elliptical or rectangular. Thus, if a substantially rectangular or elliptical exhaust nozzle were placed under a flat wing or fuselage, or adjacent to another engine, there would be substantially fewer voids to create afterbody pressure drag.
Both analytical studies and scale model wind tunnel performance tests have shown that aerodynamically blended asymmetric exhaust systems can provide significant reduction in nozzle/nacelle drag relative to round nozzles, as much as 10 percent of total aircraft drag. This is accomplished by providing a more gradual overall nacelle area closure rate, even at dry power conditions, and by providing a better aerodynamic blend of the exhaust system with the airframe nacelle, thereby eliminating locally severe area distributions. In particular, the space between adjacent nozzles of closely spaced engines is minimized, thereby avoiding one source of severe pressure drag.
Asymmetric nozzles, however, because of their noncircular cross section, create difficult mechanical design challenges due to large structural bending moments, stresses and the large actuation forces where area variation is incorporated. Since nozzle area variation is desirable for efficient operation of a gas turbine engine throughout the aircraft flight regime and is required for afterburning operation, it is advantageous to provide the asymmetric nozzle with this capability. Furthermore, it is desirable to provide a variable area nozzle having a minimum potential for flow leakage since a loss of exhaust flow through nozzle flaps or seals results in an equivalent loss in engine thrust.
To satisfy these requirements it becomes convenient to displace or rotate one portion of an asymmetric nozzle with respect to the remainder such that essentially one wall (herein denominated as a flap) of the rotated portion controls the area variation of the nozzle. Due to the large area of this flap, large actuation forces are required to move it against the exhaust pressure, thereby necessitating heavy actuators and associated hardware such as power supplies (pumps) and load-bearing structure. Since weight is a primary consideration in designing any aircraft component, it becomes desirable to reduce, and if possible to eliminate, the weight of an asymmetric exhaust nozzle area variation mechanism.
Furthermore, it is desirable to provide an asymmetric exhaust nozzle which possesses high internal aerodynamic performance in conjunction with low afterbody drag and variable geometry characteristics. Several prior attempts at providing area variation in an asymmetric nozzle have resulted in configurations wherein discontinuities in the internal aerodynamic flow path were present in one or more of the nozzle operative modes. It can be appreciated that such loss-creating mechanisms in the internal flow path, in addition to the flow leakage hereinbefore mentioned, can offset the substantial benefits obtainable through asymmetric contouring of the external profile.